US9238977B2 - Turbine shroud mounting and sealing arrangement - Google Patents

Turbine shroud mounting and sealing arrangement Download PDF

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Publication number
US9238977B2
US9238977B2 US13/683,813 US201213683813A US9238977B2 US 9238977 B2 US9238977 B2 US 9238977B2 US 201213683813 A US201213683813 A US 201213683813A US 9238977 B2 US9238977 B2 US 9238977B2
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United States
Prior art keywords
shroud
turbine
leg
casing
seal member
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Active, expires
Application number
US13/683,813
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English (en)
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US20140140833A1 (en
Inventor
Joseph Charles Albers
Robert Proctor
Monty Lee Shelton
Richard Russo, JR.
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
General Electric Co
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General Electric Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Priority to US13/683,813 priority Critical patent/US9238977B2/en
Application filed by General Electric Co filed Critical General Electric Co
Priority to CN201380060905.1A priority patent/CN104797784B/zh
Priority to CA2891616A priority patent/CA2891616A1/en
Priority to EP13818856.0A priority patent/EP2923041A1/en
Priority to BR112015010425A priority patent/BR112015010425A2/pt
Priority to PCT/US2013/064916 priority patent/WO2014081517A1/en
Priority to JP2015543049A priority patent/JP2015535565A/ja
Publication of US20140140833A1 publication Critical patent/US20140140833A1/en
Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SHELTON, MONTY LEE, ALBERS, JOSEPH CHARLES, RUSSO, Richard, Jr., PROCTOR, ROBERT
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Publication of US9238977B2 publication Critical patent/US9238977B2/en
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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/28—Supporting or mounting arrangements, e.g. for turbine casing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/127—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with a deformable or crushable structure, e.g. honeycomb
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding

Definitions

  • This invention relates generally to gas turbine engine turbines and more particularly to apparatus for sealing turbine sections of such engines.
  • a gas turbine engine includes a turbomachinery core having a high pressure compressor, a combustor, and a high pressure turbine in serial flow relationship.
  • the core is operable in a known manner to generate a primary gas flow.
  • the core exhaust gas is directed through an exhaust nozzle to generate thrust.
  • a turbofan engine uses a low pressure turbine downstream of the core to extract energy from the primary flow to drive a fan which generates propulsive thrust.
  • the low pressure turbine includes annular arrays of stationary vanes or nozzles that direct the gases exiting the combustor into rotating blades or buckets. Collectively one row of nozzles and one row of blades make up a “stage”. Typically two or more stages are used in serial flow relationship.
  • a turbine shroud apparatus for a gas turbine engine having a centerline axis includes: a shroud segment having: an arcuate body extending axially between forward and aft ends and laterally between opposed end faces, wherein each of the end faces includes seal slots formed therein; and an arcuate stationary seal member mounted to the body; a turbine vane disposed axially aft of the shroud segment; and a casing surrounding the shroud segment and the turbine vane; wherein the turbine vane is mounted to the case so as to bear against the stationary seal member, compressing it and forcing the shroud segment radially outward against the casing.
  • a turbine shroud apparatus for a gas turbine engine having a centerline axis includes: an annular array of rotatable turbine blades, each blade having an annular seal tooth projecting radially outward therefrom; a shroud surrounding the turbine blades, the shroud comprising an annular array of side-by-side shroud segments, each shroud segment having: an arcuate body extending axially between forward and aft ends and laterally between opposed end faces, wherein each of the end faces includes seal slots formed therein; and an arcuate stationary seal member mounted to the body, wherein the end faces of adjacent shroud segments abut each other and at least one spline seal is received in the seal slots so as to span the gap between adjacent shroud segments; an annular array of airfoil-shaped turbine vanes disposed axially aft of the shroud; and a casing surrounding the shroud segments and the turbine vanes; wherein each of the turbine vanes is mounted to
  • FIG. 1 a schematic cross-sectional view of a gas turbine engine constructed in accordance with the present invention
  • FIG. 2 is an enlarged view of a portion of a turbine section of the engine shown in FIG. 1 ;
  • FIG. 3 is a front elevational view of a turbine shroud segment shown in FIG. 2 ;
  • FIG. 4 is a side view of a portion of the shroud segment shown in FIG. 2 ;
  • FIG. 5 is a cross-sectional view of a portion of two side-by-side shroud segments, showing a spline seal installed therein.
  • FIGS. 1 and 2 depict a portion of a gas turbine 10 engine having, among other structures, a fan 12 , a low-pressure compressor or “booster” 14 , a high-pressure compressor 16 , a combustor 18 , a high-pressure turbine 20 , and a low-pressure turbine 22 .
  • the high-pressure compressor 16 provides compressed air that passes primarily into the combustor 18 to support combustion and partially around the combustor 18 where it is used to cool both the combustor liners and turbomachinery further downstream. Fuel is introduced into the forward end of the combustor 18 and is mixed with the air in a conventional fashion.
  • the resulting fuel-air mixture is ignited for generating hot combustion gases.
  • the hot combustion gases are discharged to the high pressure turbine 20 where they are expanded so that energy is extracted.
  • the high pressure turbine 20 drives the high-pressure compressor 16 through an outer shaft 24 .
  • the gases exiting the high-pressure turbine 20 are discharged to the low-pressure turbine 22 where they are further expanded and energy is extracted to drive the booster 14 and fan 12 through an inner shaft 26 .
  • the engine is a turbofan engine.
  • turbofan engine the principles described herein are equally applicable to turboprop, turbojet, and turbofan engines, as well as turbine engines used for other vehicles or in stationary applications.
  • the low pressure turbine 22 includes a rotor carrying a array of airfoil-shaped turbine blades 28 extending outwardly from a disk that rotates about a centerline axis “A” of the engine 10 .
  • the tip 30 of each blade 28 has one or more annular, flange-like seal teeth 32 extending radially outward therefrom.
  • a plurality of shroud segments 34 are arranged in an annulus so as to closely surround the turbine blades 28 and thereby define the outer radial flowpath boundary for the hot gas stream flowing through the rotor.
  • Each shroud segment 34 includes an arcuate body 36 extending between end faces 38 (see FIG. 3 ) and having forward and aft ends 40 and 42 . From rear to front the body 36 includes a first leg 44 which extends at an acute angle to the centerline axis A, a second leg 46 which also extends at an acute angle to the centerline axis A, a third leg 48 extending generally radially inward from the second leg 46 , and a fourth leg 50 extending generally axially forward from the third leg 48 .
  • the first leg 44 and the second leg 46 meet in a shallow “V” angle with the apex of the V facing radially outwards.
  • a boss 54 is disposed adjacent the intersection of the first and second legs 44 and 46 and includes a radially-outward-facing groove 56 formed therein.
  • each of the legs 44 , 46 , 48 , and 50 includes a slot 58 sized and shaped to receive a conventional spline seal 59 (seen in FIG. 5 ).
  • a spline seal takes the form of a thin strip of metal or other suitable material which is inserted in slots 58 . The spline seals span the gaps between shroud segments 34 .
  • a stationary seal member 60 is mounted to the radially inner face of the body 36 .
  • the seal member 60 serves the purpose of forming a non-contact rotating seal in conjunction with the seal teeth 32 .
  • the seal member 60 is configured so as to be sacrificial in the even of contact with the seal tooth 32 during operation, an event known as a “rub”.
  • Various types of sacrificial materials exist, such as nonmetallic abradable materials and honeycomb structures.
  • the seal member 60 comprises a known type of metallic honeycomb structure comprising a plurality of side-by-side cells, extending in the radial direction.
  • the seal member 60 has a back surface which conforms to the inner surface of the body 36 . It also includes a flowpath surface 62 .
  • the flowpath surface 62 comprises a plurality of cylindrical sections that define a stepped profile, with the surface of each “step” being selected to provide a desired clearance to the adjacent seal tooth 32 .
  • the seal member 60 extends radially inward beyond the first leg 44 of the body 36 , so as to create a slight interference fit, as described in more detail below.
  • the height “H” of the overhang is shown in FIG. 4 , greatly exaggerated for illustrative purposes.
  • a nozzle is positioned downstream of the rotor, and comprises a plurality of circumferentially spaced airfoil-shaped vanes 64 , each of which terminates at an arcuate tip shroud 66 .
  • Arcuate forward and aft hooks 68 and 70 extend outward from the tip shroud 66 .
  • the forward hook 68 extends axially forward and radially outward, and includes a flange 72 extending axially forward at its distal end.
  • An annular casing 74 surrounds shroud segments 34 and the vanes 64 .
  • the casing 74 includes an annular mounting slot 76 which faces axially aft, and also an annular mounting hook 78 with an L-shaped cross-sectional shape.
  • the forward flange 52 of the shroud segment 34 is received in the mounting slot 76 .
  • the slot 56 of the boss 54 receives the mounting hook 78 .
  • the forward hook 68 of the vane 64 is received in a slot defined by the mounting hook 78 .
  • the tip shroud 66 of the vane 64 bears radially outward against the shroud segment 34 .
  • the radial distance between the mounting hook 78 and the tip shroud 66 is selected such that the tip shroud 66 creates a slight interference fit with the stationary seal member 60 .
  • the seal member 60 compresses to accommodate this interference, creating a reliable seal against air leakage and holding the shroud segment 34 firmly against the mounting hook 78 .
  • a technical advantage of this configuration is a reduction in leakage through the gaps and a reduction in air temperature in the cavity.
  • the reduction in leakage and air temperature through the gaps will allow for better performance.
  • the reduction of air temperature in the cavity will help protect the case hooks from increased temperature and prevent cracking.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US13/683,813 2012-11-21 2012-11-21 Turbine shroud mounting and sealing arrangement Active 2033-11-29 US9238977B2 (en)

Priority Applications (7)

Application Number Priority Date Filing Date Title
US13/683,813 US9238977B2 (en) 2012-11-21 2012-11-21 Turbine shroud mounting and sealing arrangement
CA2891616A CA2891616A1 (en) 2012-11-21 2013-10-15 Turbine shroud mounting and sealing arrangement
EP13818856.0A EP2923041A1 (en) 2012-11-21 2013-10-15 Turbine shroud mounting and sealing arrangement
BR112015010425A BR112015010425A2 (pt) 2012-11-21 2013-10-15 aparelhos de invólucro de turbina para um motor de turbina a gás
CN201380060905.1A CN104797784B (zh) 2012-11-21 2013-10-15 涡轮机护罩安装和密封结构
PCT/US2013/064916 WO2014081517A1 (en) 2012-11-21 2013-10-15 Turbine shroud mounting and sealing arrangement
JP2015543049A JP2015535565A (ja) 2012-11-21 2013-10-15 タービンシュラウドの取り付け及び封止の構成

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/683,813 US9238977B2 (en) 2012-11-21 2012-11-21 Turbine shroud mounting and sealing arrangement

Publications (2)

Publication Number Publication Date
US20140140833A1 US20140140833A1 (en) 2014-05-22
US9238977B2 true US9238977B2 (en) 2016-01-19

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
US13/683,813 Active 2033-11-29 US9238977B2 (en) 2012-11-21 2012-11-21 Turbine shroud mounting and sealing arrangement

Country Status (7)

Country Link
US (1) US9238977B2 (zh)
EP (1) EP2923041A1 (zh)
JP (1) JP2015535565A (zh)
CN (1) CN104797784B (zh)
BR (1) BR112015010425A2 (zh)
CA (1) CA2891616A1 (zh)
WO (1) WO2014081517A1 (zh)

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20180340437A1 (en) * 2017-02-24 2018-11-29 General Electric Company Spline for a turbine engine
US20180355754A1 (en) * 2017-02-24 2018-12-13 General Electric Company Spline for a turbine engine
US20180355755A1 (en) * 2017-02-24 2018-12-13 General Electric Company Spline for a turbine engine
US20180355753A1 (en) * 2017-02-24 2018-12-13 General Electric Company Spline for a turbine engine
US20190085713A1 (en) * 2017-09-21 2019-03-21 Safran Aircraft Engines Turbine sealing assembly for turbomachinery
US10494940B2 (en) * 2016-04-05 2019-12-03 MTU Aero Engines AG Seal segment assembly including mating connection for a turbomachine
US10494946B2 (en) 2013-03-14 2019-12-03 General Electric Company Method of making a turbine shroud
US20200063586A1 (en) * 2018-08-24 2020-02-27 General Electric Company Spline Seal with Cooling Features for Turbine Engines
US11608752B2 (en) 2021-02-22 2023-03-21 General Electric Company Sealing apparatus for an axial flow turbomachine
US20230175412A1 (en) * 2019-09-13 2023-06-08 Safran Aircraft Engines Turbomachine sealing ring

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ES2620482T3 (es) * 2012-08-09 2017-06-28 MTU Aero Engines AG Impermeabilización del canal de flujo de una turbomáquina
US10612407B2 (en) 2013-02-28 2020-04-07 United Technologies Corporation Contoured blade outer air seal for a gas turbine engine
JP6233578B2 (ja) * 2013-12-05 2017-11-22 株式会社Ihi タービン
US20180347399A1 (en) * 2017-06-01 2018-12-06 Pratt & Whitney Canada Corp. Turbine shroud with integrated heat shield
US10519790B2 (en) * 2017-06-15 2019-12-31 General Electric Company Turbine shroud assembly
US20200141276A1 (en) * 2018-11-07 2020-05-07 General Electric Company Turbine shroud with lapped seal segments
FR3096723B1 (fr) * 2019-05-29 2022-03-25 Safran Helicopter Engines Anneau d’etancheite pour une roue de turbine de turbomachine
CN115263808B (zh) * 2022-09-28 2023-02-21 中国航发四川燃气涡轮研究院 集成式双转子航空发动机的中介机匣

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Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10494946B2 (en) 2013-03-14 2019-12-03 General Electric Company Method of making a turbine shroud
US10494940B2 (en) * 2016-04-05 2019-12-03 MTU Aero Engines AG Seal segment assembly including mating connection for a turbomachine
US20180355753A1 (en) * 2017-02-24 2018-12-13 General Electric Company Spline for a turbine engine
US20180340437A1 (en) * 2017-02-24 2018-11-29 General Electric Company Spline for a turbine engine
US20180355755A1 (en) * 2017-02-24 2018-12-13 General Electric Company Spline for a turbine engine
US20180355754A1 (en) * 2017-02-24 2018-12-13 General Electric Company Spline for a turbine engine
US10648362B2 (en) * 2017-02-24 2020-05-12 General Electric Company Spline for a turbine engine
US10655495B2 (en) * 2017-02-24 2020-05-19 General Electric Company Spline for a turbine engine
US20190085713A1 (en) * 2017-09-21 2019-03-21 Safran Aircraft Engines Turbine sealing assembly for turbomachinery
US10871079B2 (en) * 2017-09-21 2020-12-22 Safran Aircraft Engines Turbine sealing assembly for turbomachinery
US20200063586A1 (en) * 2018-08-24 2020-02-27 General Electric Company Spline Seal with Cooling Features for Turbine Engines
US10982559B2 (en) * 2018-08-24 2021-04-20 General Electric Company Spline seal with cooling features for turbine engines
US20230175412A1 (en) * 2019-09-13 2023-06-08 Safran Aircraft Engines Turbomachine sealing ring
US11952901B2 (en) * 2019-09-13 2024-04-09 Safran Aircraft Engines Turbomachine sealing ring
US11608752B2 (en) 2021-02-22 2023-03-21 General Electric Company Sealing apparatus for an axial flow turbomachine

Also Published As

Publication number Publication date
CA2891616A1 (en) 2014-05-30
CN104797784A (zh) 2015-07-22
JP2015535565A (ja) 2015-12-14
CN104797784B (zh) 2016-09-14
WO2014081517A1 (en) 2014-05-30
BR112015010425A2 (pt) 2018-04-10
EP2923041A1 (en) 2015-09-30
US20140140833A1 (en) 2014-05-22

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